US20230340906A1 - Counter-rotating turbine - Google Patents
Counter-rotating turbine Download PDFInfo
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- US20230340906A1 US20230340906A1 US17/713,649 US202217713649A US2023340906A1 US 20230340906 A1 US20230340906 A1 US 20230340906A1 US 202217713649 A US202217713649 A US 202217713649A US 2023340906 A1 US2023340906 A1 US 2023340906A1
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Images
Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/06—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages
- F02C3/067—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor the compressor comprising only axial stages having counter-rotating rotors
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C3/00—Gas-turbine plants characterised by the use of combustion products as the working fluid
- F02C3/04—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor
- F02C3/107—Gas-turbine plants characterised by the use of combustion products as the working fluid having a turbine driving a compressor with two or more rotors connected by power transmission
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/12—Cooling of plants
- F02C7/14—Cooling of plants of fluids in the plant, e.g. lubricant or fuel
- F02C7/141—Cooling of plants of fluids in the plant, e.g. lubricant or fuel of working fluid
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
- F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
- F02C7/36—Power transmission arrangements between the different shafts of the gas turbine plant, or between the gas-turbine plant and the power user
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
- F02K—JET-PROPULSION PLANTS
- F02K3/00—Plants including a gas turbine driving a compressor or a ducted fan
- F02K3/02—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber
- F02K3/04—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type
- F02K3/06—Plants including a gas turbine driving a compressor or a ducted fan in which part of the working fluid by-passes the turbine and combustion chamber the plant including ducted fans, i.e. fans with high volume, low pressure outputs, for augmenting the jet thrust, e.g. of double-flow type with front fan
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2220/00—Application
- F05D2220/60—Application making use of surplus or waste energy
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/20—Heat transfer, e.g. cooling
- F05D2260/213—Heat transfer, e.g. cooling by the provision of a heat exchanger within the cooling circuit
-
- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
- F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
- F05D2260/00—Function
- F05D2260/40—Transmission of power
- F05D2260/403—Transmission of power through the shape of the drive components
- F05D2260/4031—Transmission of power through the shape of the drive components as in toothed gearing
- F05D2260/40311—Transmission of power through the shape of the drive components as in toothed gearing of the epicyclical, planetary or differential type
Definitions
- the present subject matter relates generally to a counter-rotating turbine of a gas turbine engine. More specifically, the subject matter relates to a counter-rotating low pressure turbine of a gas turbine engine.
- a volume of air 58 enters the gas turbine engine 10 through an associated inlet 60 of the outer nacelle 50 and/or the fan section 14 .
- a first portion of air 62 from the volume of air 58 is directed or routed into the bypass airflow passage 56 and a second portion of air 64 from the volume of air 58 is directed or routed into the LP compressor 22 .
- the ratio between the first portion of air 62 and the second portion of air 64 is commonly known as a bypass ratio.
- the pressure of the second portion of air 64 is then increased as it is routed through the HP compressor 24 and into the combustion section 26 , where it is mixed with fuel and burned to provide combustion gases 66 .
- the gas turbine engine 10 may instead be configured as, e.g., a turbojet engine, a turboshaft engine, a turboprop engine, etc., and further may be configured as an aeroderivative gas turbine engine or an industrial gas turbine engine.
- FIG. 2 a simplified, schematic view of a gas turbine engine 10 in accordance with an exemplary aspect of the present disclosure is provided.
- the exemplary gas turbine engine 10 depicted in FIG. 2 may be configured in substantially the same manner as exemplary gas turbine engine 10 described above with reference to FIG. 1 .
- the exemplary gas turbine engine 10 includes a waste heat recovery system 86 .
- the exemplary waste heat recovery system 86 is generally configured to extract heat from a heat source (e.g., a heat source not fully utilizing the heat being extracted therefrom) and transfer such extracted heat to a heat sink, such that the heat sink may more efficiently utilize such extracted heat.
- a heat source e.g., a heat source not fully utilizing the heat being extracted therefrom
- the waste heat recovery system 86 generally includes a heat source exchanger 250 (i.e., a heat exchanger configured to extract heat for the waste heat recovery system 86 from a heat source of the gas turbine engine 10 ), a heat sink exchanger 96 (i.e., a heat exchanger configured to transfer heat from the waste heat recovery system 86 to a heat sink of the gas turbine engine 10 ), a thermal transfer bus 98 , and a pump 100 .
- a heat source exchanger 250 i.e., a heat exchanger configured to extract heat for the waste heat recovery system 86 from a heat source of the gas turbine engine 10
- a heat sink exchanger 96 i.e., a heat exchanger configured to transfer heat from the waste heat recovery system 86 to a heat sink of the gas turbine engine 10
- a thermal transfer bus 98 i.e., a heat exchanger configured to transfer heat from the waste heat recovery system 86 to a heat sink of the gas turbine engine 10
- a thermal transfer bus 98 i.
- the heat source exchanger 250 is in thermal communication with the exhaust section 32 of the gas turbine engine 10 such that the heat source exchanger 250 extracts heat from the exhaust section 32 .
- the heat source exchanger 250 may be integrated into a strut extending through the exhaust section 32 or a liner defining at least in part the exhaust section 32 , or alternatively may be positioned at any other suitable location in thermal communication with an airflow/ gases through the exhaust section 32 of the gas turbine engine 10 .
- the term “proximate the downstream end,” with reference to the HP compressor 24 refers to a location closer to an exit 102 of the HP compressor 24 than an inlet 104 to the HP compressor 24 and upstream of the combustion chamber 92 .
- the heat sink exchanger 96 may be integrated into, or coupled to, a strut or guide vane, such as a diffuser, positioned at the exit 102 of the HP compressor 24 and upstream of the combustion chamber 92 of the combustion section 26 .
- the heat sink exchanger 96 may be integrated into, or coupled to, one or more fuel nozzles 90 of the fuel delivery system 80 .
- the heat sink exchanger 96 is disposed in a portion of the gas turbine engine 10 upstream of the combustion section 26 .
- the heat sink exchanger 96 may be disposed within the fuel delivery system 80 upstream of the combustion section 26 , within the core air flowpath 37 upstream of the combustion section 26 , or other locations upstream of the combustion section 26 .
- the term “upstream” refers to the direction of flow and not the axial direction.
- the heat sink exchanger 96 may be forward of, aft of, or in line with the combustion section 26 in the axial direction A.
- the fuel delivery system 80 may deliver conventional fuel or, alternatively or additionally, may deliver H 2 fuel.
- the gas turbine engine further includes a waste heat recovery system 86 with a thermal transfer bus 98 having a thermal transfer fluid configured to flow therethrough.
- the heat source exchanger 250 is part of the waste heat recovery system 86 and is coupled to the thermal transfer bus 98 , such that the thermal transfer fluid passing through the thermal transfer bus 98 passes through the heat source exchanger 250 to extract heat from the exhaust section 32 of the gas turbine engine 10 .
- the heat source exchanger 250 is a frame-integrated heat exchanger that is integrated into the frame of the turbomachine 16 . According to one or more embodiments, the heat source exchanger 250 is a heat exchanger formed separately from the frame of the turbomachine 16 .
- FIGS. 1 and 3 show two stages of turbine rotor blades with a single stage of turbine guide vanes therebetween upstream of the final stage LP turbine rotor blades 241
- one or more embodiments may include additional stages of turbine rotor blades and turbine guide vanes.
- the LP turbine 30 may further include third turbine guide vanes and third stage LP turbine rotor blades in serial order downstream of the second stage LP turbine rotor blades 223 and may further include fourth turbine guide vanes and fourth stage LP turbine rotor blades in serial order downstream of the third stage LP turbine rotor blades, and so on.
- the gearbox 260 is further coupled to the LP shaft or spool 36 , such that the first stage LP turbine rotor blades 221 , the second stage LP turbine rotor blades 223 , and the final stage LP turbine rotor blades 241 drive the LP shaft or spool 36 through the gearbox 260 .
- the LP turbine 30 supports operation of the LP compressor 22 and/or rotation of the fan 38 (see FIG. 1 ).
- the gearbox 260 is disposed partially or entirely on an inner side of the LP turbine 30 in the radial direction R.
- the gearbox 260 may be structured as a planetary gear system in which that the LP shaft or spool 36 is fixed to or connected to a sun gear 261 , the second LP turbine spool 245 is connected to planet gears 263 that are disposed around the sun gear 261 and are meshed with the sun gear 261 , and the first LP turbine spool 225 is fixed to or connected to a ring gear 265 disposed around the planet gears 263 and is meshed with the planet gears 263 .
- first LP turbine spool 225 and the LP shaft or spool 36 rotate in the first direction around the longitudinal centerline 12 , with the first LP turbine spool 225 rotating at a higher speed than the LP shaft or spool 36 , and the second LP turbine spool 245 rotates in the second direction at a lower speed than the first LP turbine spool 225 .
- the final stage LP turbine rotor blades 241 have a significantly greater height in the radial direction R than the first stage LP turbine rotor blades 221 and the second stage LP turbine rotor blades 223 .
- the final stage LP turbine rotor blades 241 also have a significantly greater height in the radial direction R than final stage LP turbine rotor blades of conventional high speed low pressure turbines that rotate at the same speed in the same direction as the preceding stages of LP turbine rotor blades.
- the first stage LP turbine rotor blades 221 , the second stage LP turbine rotor blades 223 , and the final stage LP turbine rotor blades 241 are formed of nickel alloys, or are formed of a material that comprises nickel alloys. According to one or more embodiments, the final stage LP turbine rotor blades 241 are formed of a different material from the first stage LP turbine rotor blades 221 and the second stage LP turbine rotor blades 223 . According to one or more embodiments, the final stage LP turbine rotor blades 241 are formed of titanium aluminide or a material comprising titanium aluminide.
- the LP turbine 30 is structured such that combustion gases exits the LP turbine 30 and enters the exhaust section 32 at a Mach number of 1 ⁇ 4 Ma or less.
- a traditional heat exchanger separate from the frame of the turbomachine 16 may be employed as the heat source exchanger 250 in the exhaust section 32 .
- the turbine section and the exhaust section of one or more of these clauses further comprising a heat source exchanger disposed downstream of the final stage LP turbine blades.
- a turbine section and an exhaust section of a gas turbine engine comprises a low pressure (LP) turbine comprising a plurality of stages of LP turbine blades that rotate in a first direction at a first speed, and final stage LP turbine blades downstream of the plurality of stages of LP turbine blades that rotate in a second direction opposite the first direction at a second speed, wherein the second speed is lower than the first speed.
- LP low pressure
- a gas turbine engine comprises a low pressure (LP) turbine comprising a plurality of stages of LP turbine blades that rotate in a first direction at a first speed, and final stage LP turbine blades downstream of the plurality of stages of LP turbine blades that rotate in a second direction opposite the first direction at a second speed, wherein the second speed is lower than the first speed.
- LP low pressure
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Abstract
Description
- The present subject matter relates generally to a counter-rotating turbine of a gas turbine engine. More specifically, the subject matter relates to a counter-rotating low pressure turbine of a gas turbine engine.
- A gas turbine engine typically includes a fan and a turbomachine. The turbomachine generally includes an inlet, one or more compressors, a combustor, and one or more turbines. The compressors compress air channeled to the combustor where the compressed air is mixed with fuel. The mixture is then ignited for generating hot combustion gases. The combustion gases are channeled to the turbine(s) that extract energy from the combustion gases for powering the compressor(s), and for producing useful work to propel an aircraft in flight and/or to power a load, such as an electrical generator.
- During operation of the gas turbine engine, a relatively large amount of heat energy is generated through the compression process by the compressors and through the combustion process within the combustor. While a substantial portion of the heat energy is extracted through the one or more turbines, a portion of such heat energy is exhausted to atmosphere. Such may lead to a loss of efficiency of the gas turbine engine. Accordingly, a system and/or a method for operating a gas turbine engine in a manner to increase an efficiency of the gas turbine engine, and to reduce an amount of waste heat exhausted to atmosphere, would be useful.
- A full and enabling disclosure, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:
-
FIG. 1 is a schematic cross-sectional view of a gas turbine engine according to one or more embodiments. -
FIG. 2 is a simplified schematic view of a gas turbine engine and a waste heat recovery system in accordance with one or more embodiments. -
FIG. 3 is a schematic view of a portion of a turbine section and an exhaust section of a gas turbine engine according to one or more embodiments. - Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present disclosure.
- Reference now will be made in detail to embodiments of the disclosure, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the disclosure, and is not intend to be limiting in any way. In fact, it will be apparent to those skilled in the art that various modifications and variations may be made in the present disclosure without departing from the scope or the spirit of the disclosure. For instance, features illustrated or described as part of one embodiment may be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present disclosure covers such modifications and variations as come within the scope of the appended claims and their equivalents.
- As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.
- The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.
- The terms “coupled,” “fixed,” “attached to,” and the like refer to both direct coupling, fixing, or attaching, as well as indirect coupling, fixing, or attaching through one or more intermediate components or features, unless otherwise specified herein.
- The singular forms “a”, “an”, and “the” include plural references unless the context clearly dictates otherwise.
- Approximating language, as used herein throughout the specification and claims, is applied to modify any quantitative representation that could permissibly vary without resulting in a change in the basic function to which it is related. Accordingly, a value modified by a term or terms, such as “about”, “approximately”, and “substantially”, are not to be limited to the precise value specified. In at least some instances, the approximating language may correspond to the precision of an instrument for measuring the value, or the precision of the methods or the machines for constructing or manufacturing the components and/or systems. For example, the approximating language may refer to being within a 1, 2, 4, 10, 15, or 20 percent margin. These approximating margins may apply to a single value, either or both endpoints defining numerical ranges, and/or the margin for ranges between endpoints.
- Here and throughout the specification and claims, range limitations are combined and interchanged, such ranges are identified and include all the sub-ranges contained therein unless context or language indicates otherwise. For example, all ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other.
- The term “turbomachine” or “turbomachinery” refers to a machine including one or more compressors, a heat generating section (e.g., a combustion section), and one or more turbines that together generate a torque output.
- The term “gas turbine engine” refers to an engine having a turbomachine as all or a portion of its power source. Example gas turbine engines include turbofan engines, turboprop engines, turbojet engines, turboshaft engines, etc., as well as hybrid-electric versions of one or more of these engines.
- The term “combustion section” refers to any heat addition system for a turbomachine. For example, the term combustion section may refer to a section including one or more of a deflagrative combustion assembly, a rotating detonation combustion assembly, a pulse detonation combustion assembly, or other appropriate heat addition assembly. In certain example embodiments, the combustion section may include an annular combustor, a can combustor, a cannular combustor, a trapped vortex combustor (TVC), or other appropriate combustion system, or combinations thereof.
- The terms “low” and “high”, or their respective comparative degrees (e.g., -er, where applicable), when used with a compressor, a turbine, a shaft, or spool components, etc., each refers to relative speeds within an engine unless otherwise specified. For example, a “low turbine” or “low speed turbine” defines a component configured to operate at a rotational speed, such as a maximum allowable rotational speed, lower than a “high turbine” or a “high speed turbine” of the engine.
- In a conventional gas turbine engine, a significant amount of heat is exhausted to atmosphere as waste heat. One way of reducing the amount of waste heat and improving an efficiency of the gas turbine engine is recovering some of the waste heat via a heat source exchanger positioned within an exhaust section of the gas turbine engine.
- While conventional gas turbine engines have low pressure turbines that operate at comparatively low speeds to drive the low speed compressor and/or the fan, gas turbine engines may instead employ high speed low pressure turbines that rotate at a higher speed than the conventional low pressure turbines. The high speed low pressure turbines may include a reducing gearbox in the fan module such that the rotation speed of the fan is reduced from the high rotation speed of the high speed low pressure turbines.
- However, due to the increased rotation speeds, the blades of the high speed low pressure turbines experience increased stresses compared to conventional low pressure turbines. Because an increase in size of the blades also increases an amount of stress experienced thereby, a material capability of the blades may require a size of the blades to be reduced to compensate for the increased stresses caused by the higher rotation speeds. Therefore, high speed low pressure turbines may have a limited exhaust annulus area due to blade material capability.
- Additionally, the higher speeds of the blades within the high speed low pressure turbines may also result in a high exhaust Mach number. That is, the combustion gases exiting the high speed low pressure turbines and entering the exhaust section may have a high Mach number. The high Mach number entering the exhaust section may cause losses in waste heat recovery for heat source exchangers within the exhaust section. This loss is proportional to a square of the Mach number of the combustion gases entering the heat source exchangers.
- A counter-rotating turbine according to one or more embodiments allows for larger turbine blades within high speed low pressure turbines and a more axially compact package, while reducing the exhaust Mach number and significantly reducing losses in waste heat recovery for the heat source exchangers within the exhaust section. According to one or more embodiments, by allowing for lower speed final stage low pressure turbine blades, the final stage low pressure turbine blades may be formed to have larger radial dimensions and thus larger annulus flow areas, which in turn reduces a Mach number of the flow entering the heat source exchanger downstream of the final stage low pressure turbine blades.
- Referring now to the drawings, wherein identical numerals indicate the same elements throughout the figures,
FIG. 1 is a schematic cross-sectional view of a gas turbine engine in accordance with an exemplary embodiment of the present disclosure. More particularly, for the embodiment ofFIG. 1 , the gas turbine engine is a high-bypassturbofan jet engine 10, referred to herein as “gas turbine engine 10.” As shown inFIG. 1 , thegas turbine engine 10 defines an axial direction A (extending parallel to alongitudinal centerline 12 provided for reference), a radial direction R, and a circumferential direction (i.e., a direction extending about the axial direction A; not depicted). In general, thegas turbine engine 10 includes afan section 14 and aturbomachine 16 disposed downstream from thefan section 14. - The
exemplary turbomachine 16 depicted generally includes a substantially tubularouter casing 18 that defines anannular inlet 20. Theouter casing 18 encases, in serial flow relationship, a compressor section including a booster or low pressure (LP)compressor 22 and a high pressure (HP)compressor 24, acombustion section 26, a turbine section including a high pressure (HP)turbine 28 and a low pressure (LP)turbine 30, and anexhaust section 32. The compressor section, thecombustion section 26, and the turbine section together define acore air flowpath 37 extending from theannular inlet 20 through theLP compressor 22, the HPcompressor 24, thecombustion section 26, the HPturbine 28, theLP turbine 30 and theexhaust section 32. A high pressure (HP) shaft or aspool 34 drivingly connects theHP turbine 28 to theHP compressor 24. A low pressure (LP) shaft or aspool 36 drivingly connects theLP turbine 30 to theLP compressor 22. - For the embodiment depicted, the
fan section 14 includes afan 38 having a plurality offan blades 40 coupled to adisk 42 in a spaced apart manner. As depicted, thefan blades 40 extend outwardly fromdisk 42 generally along the radial direction R. Thefan blades 40 anddisk 42 are together rotatable about thelongitudinal centerline 12 by the LP shaft or thespool 36. - Referring still to the exemplary embodiment of
FIG. 1 , thedisk 42 is covered byrotatable spinner cone 48 aerodynamically contoured to promote an airflow through the plurality offan blades 40. Additionally, theexemplary fan section 14 includes an annular fan casing orouter nacelle 50 that circumferentially surrounds thefan 38 and/or at least a portion of theturbomachine 16. For the embodiment depicted, theouter nacelle 50 is supported relative to theturbomachine 16 by a plurality of circumferentially-spaced outlet guide vanes 52. Moreover, adownstream section 54 of theouter nacelle 50 extends over an outer portion ofturbomachine 16 so as to define abypass airflow passage 56 therebetween. - During operation of the
gas turbine engine 10, a volume ofair 58 enters thegas turbine engine 10 through an associatedinlet 60 of theouter nacelle 50 and/or thefan section 14. As the volume ofair 58 passes across thefan blades 40, a first portion ofair 62 from the volume ofair 58 is directed or routed into thebypass airflow passage 56 and a second portion ofair 64 from the volume ofair 58 is directed or routed into theLP compressor 22. The ratio between the first portion ofair 62 and the second portion ofair 64 is commonly known as a bypass ratio. The pressure of the second portion ofair 64 is then increased as it is routed through theHP compressor 24 and into thecombustion section 26, where it is mixed with fuel and burned to providecombustion gases 66. - The
combustion gases 66 are routed through theHP turbine 28 where a portion of thermal and/or kinetic energy from thecombustion gases 66 is extracted via sequential stages of HPturbine stator vanes 68 that are coupled to an inner casing (not shown) and HPturbine rotor blades 70 that are coupled to the HP shaft orspool 34, thus causing the HP shaft orspool 34 to rotate, thereby supporting operation of theHP compressor 24. - The
combustion gases 66 are then routed through theLP turbine 30 where a second portion of thermal and kinetic energy is extracted from thecombustion gases 66 via first stage LPturbine rotor blades 221 and second stage LPturbine rotor blades 223 that rotate together in a first direction around thelongitudinal centerline 12, and final stage LPturbine rotor blades 241 that rotate in a second direction around thelongitudinal centerline 12 opposite the first direction. Firstturbine guide vanes 210 are disposed upstream of the first stage LPturbine rotor blades 221, and secondturbine guide vanes 230 are disposed between the first stage LPturbine rotor blades 221 and the second stage LPturbine rotor blades 223. The first stage LPturbine rotor blades 221 and the second stage LPturbine rotor blades 223 are connected via a firstLP turbine spool 225 that rotates in the first direction around thelongitudinal centerline 12 and is coupled to agearbox 260. The final stage LPturbine rotor blades 241 are connected to a secondLP turbine spool 245 that rotates in the second direction and is also coupled to thegearbox 260. Thegearbox 260 is further coupled to the LP shaft orspool 36, such that the first stage LPturbine rotor blades 221, the second stage LPturbine rotor blades 223, and the final stage LPturbine rotor blades 241 drive the LP shaft orspool 36 to rotate through thegearbox 260. Thus, theLP turbine 30 supports operation of theLP compressor 22 and/or rotation of thefan 38. - The
combustion gases 66 are subsequently routed through theexhaust section 32 of theturbomachine 16 to provide propulsive thrust. Theexhaust section 32 includes a heat source exchanger 250 disposed immediately downstream of the final stage LPturbine rotor blades 241. - Simultaneously, the pressure of the first portion of
air 62 is substantially increased as the first portion ofair 62 is routed through thebypass airflow passage 56 before it is exhausted from a fannozzle exhaust section 78 of thegas turbine engine 10, also providing propulsive thrust. TheHP turbine 28, theLP turbine 30, and theexhaust section 32 at least partially define a hot gas path for routing thecombustion gases 66 through theturbomachine 16. - The exemplary
gas turbine engine 10 depicted inFIG. 1 is by way of example only, and in other exemplary embodiments, thegas turbine engine 10 may have any other suitable configuration. For example, in other exemplary embodiments, thegas turbine engine 10 may instead be configured as any other suitable turbomachine including, e.g., any other suitable number of shafts or spools, and excluding, e.g., thefan 38 and/or including, e.g., a gearbox between thefan 38 and the LP shaft orspool 36, avariable pitch fan 38, etc. Accordingly, in other exemplary embodiments, thegas turbine engine 10 may instead be configured as, e.g., a turbojet engine, a turboshaft engine, a turboprop engine, etc., and further may be configured as an aeroderivative gas turbine engine or an industrial gas turbine engine. - Referring now to
FIG. 2 , a simplified, schematic view of agas turbine engine 10 in accordance with an exemplary aspect of the present disclosure is provided. The exemplarygas turbine engine 10 depicted inFIG. 2 may be configured in substantially the same manner as exemplarygas turbine engine 10 described above with reference toFIG. 1 . - For example, as is shown, the
gas turbine engine 10 generally includes afan section 14 and aturbomachine 16. Theturbomachine 16 includes in serial flow order a compressor section having anLP compressor 22 and anHP compressor 24, acombustion section 26, a turbine section including anHP turbine 28 and anLP turbine 30, and anexhaust section 32. The compressor section, thecombustion section 26, the turbine section, and theexhaust section 32 generally define acore air flowpath 37 extending therethrough. - A
fuel delivery system 80 is also included for providing a flow of fuel to thecombustion section 26 of thegas turbine engine 10, and more specifically to acombustion chamber 92 of thecombustion section 26. For example, thefuel delivery system 80 generally includes a plurality offuel lines 84 and afuel nozzle 90. Thefuel nozzle 90 may receive a flow of fuel from the plurality offuel lines 84 and further may receive compressed air from the compressor section (e.g., the HP compressor 24). Thefuel nozzle 90 may accordingly provide a mixture of compressed air and fuel to thecombustion chamber 92, wherein such mixture of compressed air and fuel is combusted to generate combustion gases. Thefuel delivery system 80 generally also includes afuel source 82 and apump 88 in fluid communication with the one ormore fuel lines 84, thepump 88 configured for increasing a pressure of a fuel flow from thefuel source 82 and through the one or more fuel lines 84. - Moreover, the
turbomachine 16 andfan section 14 are at least partially surrounded by anouter nacelle 50, with theturbomachine 16 supported relative to theouter nacelle 50 through a plurality of outlet guide vanes 52. Theouter nacelle 50 defines abypass airflow passage 56 with theturbomachine 16. Afirst portion 62 of an airflow from thefan section 14 is provided through theturbomachine 16 as a core airflow, and a second portion of theair 64 from thefan section 14 is provided through thebypass airflow passage 56 as a bypass airflow. - In addition, the exemplary
gas turbine engine 10 includes a wasteheat recovery system 86. The exemplary wasteheat recovery system 86 is generally configured to extract heat from a heat source (e.g., a heat source not fully utilizing the heat being extracted therefrom) and transfer such extracted heat to a heat sink, such that the heat sink may more efficiently utilize such extracted heat. - Referring particularly to the exemplary waste
heat recovery system 86 depicted inFIG. 2 , the wasteheat recovery system 86 generally includes a heat source exchanger 250 (i.e., a heat exchanger configured to extract heat for the wasteheat recovery system 86 from a heat source of the gas turbine engine 10), a heat sink exchanger 96 (i.e., a heat exchanger configured to transfer heat from the wasteheat recovery system 86 to a heat sink of the gas turbine engine 10), athermal transfer bus 98, and apump 100. Each of these components is described in greater detail as follows. - For the embodiment shown, the
heat source exchanger 250 is in thermal communication with theexhaust section 32 of thegas turbine engine 10 such that theheat source exchanger 250 extracts heat from theexhaust section 32. Theheat source exchanger 250 may be integrated into a strut extending through theexhaust section 32 or a liner defining at least in part theexhaust section 32, or alternatively may be positioned at any other suitable location in thermal communication with an airflow/ gases through theexhaust section 32 of thegas turbine engine 10. - Moreover, for the exemplary embodiment depicted, the
heat sink exchanger 96 is in thermal communication with the heat sink, which may generally be the compressor section, thefuel delivery system 80, and/or any other section of the aircraft requiring heat. For example, for the embodiment depicted inFIG. 2 , theheat sink exchanger 96 is in thermal communication with the compressor section at a location proximate a downstream end of the compressor section, or more specifically still with a location proximate a downstream end of theHP compressor 24 of the compressor section. As used herein, the term “proximate the downstream end,” with reference to the compressor section refers to a location closer to an exit of the compressor section than an inlet to the compressor section and upstream of thecombustion chamber 92. Similarly, as used herein, the term “proximate the downstream end,” with reference to theHP compressor 24 refers to a location closer to anexit 102 of theHP compressor 24 than aninlet 104 to theHP compressor 24 and upstream of thecombustion chamber 92. For example, in certain embodiments, theheat sink exchanger 96 may be integrated into, or coupled to, a strut or guide vane, such as a diffuser, positioned at theexit 102 of theHP compressor 24 and upstream of thecombustion chamber 92 of thecombustion section 26. Additionally, or alternatively, theheat sink exchanger 96 may be integrated into, or coupled to, one ormore fuel nozzles 90 of thefuel delivery system 80. - According to one or more embodiments, the
heat sink exchanger 96 is disposed in a portion of thegas turbine engine 10 upstream of thecombustion section 26. Theheat sink exchanger 96 may be disposed within thefuel delivery system 80 upstream of thecombustion section 26, within thecore air flowpath 37 upstream of thecombustion section 26, or other locations upstream of thecombustion section 26. Here, the term “upstream” refers to the direction of flow and not the axial direction. As such, theheat sink exchanger 96 may be forward of, aft of, or in line with thecombustion section 26 in the axial direction A. Thefuel delivery system 80 may deliver conventional fuel or, alternatively or additionally, may deliver H2 fuel. - For the embodiment shown, the
heat sink exchanger 96 is in thermal communication with theexit 102 of theHP compressor 24 of thegas turbine engine 10. The “compressor exit” refers to an outlet of theHP compressor 24. In such a manner, theheat sink exchanger 96 may add heat to an airflow through thecore air flowpath 37 prior to such airflow entering thecombustion chamber 92, resulting in a more efficientgas turbine engine 10. - Further, the waste
heat recovery system 86 includes thethermal transfer bus 98. Thethermal transfer bus 98 includes a thermal transfer fluid and extends from the heat source exchanger 250 to theheat sink exchanger 96. In such a manner, the thermal bus is configured to transfer the thermal transfer fluid from the heat source exchanger 250 (wherein the thermal transfer fluid has accepted heat from the airflow through, for the embodiment shown, theexhaust section 32 of the gas turbine engine 10) to the heat sink exchanger 96 (wherein the thermal transfer fluid transfers heat to the airflow through, for the embodiment shown, theexit 102 of theHP compressor 24, or other location upstream of the combustion chamber 92). Thethermal transfer bus 98 may include one or more pipes, conduits, etc. arranged in series, parallel, or some combination thereof. - In one or more embodiments, the
gas turbine engine 10 may define a relatively high overall pressure ratio. As used herein, the term overall pressure ratio refers to a ratio of a pressure of the air through theturbomachine 16 at an outlet of the compressor section (i.e., theexit 102 of theHP compressor 24 for the embodiment shown) to a pressure of the air through theturbomachine 16 at an inlet of the compressor section (i.e., aninlet 106 of theLP compressor 22 for the embodiment shown). In at least certain exemplary embodiments, the overall pressure ratio of thegas turbine engine 10 depicted inFIG. 2 may be at least about twenty-five. For example, in certain exemplary embodiments, the overall pressure ratio of thegas turbine engine 10 depicted inFIG. 2 may be at least about twenty-eight, such as at least about thirty, such as at least about thirty-two, such as up to about seventy-five. - In order to allow or facilitate the waste
heat recovery system 86 to transfer heat from the heat source exchanger 250 to theheat sink exchanger 96, the exemplary wasteheat recovery system 86 may further include thepump 100 in fluid communication with thethermal transfer bus 98 downstream of theheat source exchanger 250 and upstream of theheat sink exchanger 96 for increasing a temperature and a pressure of the thermal transfer fluid in thethermal transfer bus 98. For example, in certain exemplary embodiments, thepump 100 may be configured to provide at least about a twenty-five pounds per square inch (“psi”) pressurize rise, such as at least about a fifty psi of pressurize rise, such as at least about a one hundred psi of pressurize rise, such as up to about five thousand psi pressure rise, in the thermal transfer fluid through thethermal transfer bus 98, and similarly may be configured to provide at least about one hundred fifty (150) degrees Celsius temperature rise, such as at least about two hundred fifty (250) degrees Celsius temperature rise, and up to about one thousand (1,000) degrees Celsius temperature rise, in the thermal transfer fluid through thethermal transfer bus 98. Thepump 100 may be powered through, e.g., one or more of the shafts or spools of thegas turbine engine 10, or alternatively may be powered by an electric motor, hydraulic motor, pneumatic motor, or any other suitable power source. In other exemplary embodiments, however, thepump 100 may have any other suitable configuration. For example, in other embodiments, thepump 100 may be configured to create any other suitable temperature and/or pressure rise, or some other suitable device or configuration may be provided to increase a temperature and/or pressure of the thermal fluid through thethermal transfer bus 98 and provide for the flow of thermal fluid through thethermal transfer bus 98. - Referring still to
FIG. 2 , it will be appreciated that the exemplarythermal transfer bus 98 is a closed loopthermal transfer bus 98 further extending from theheat sink exchanger 96 back to theheat source exchanger 250. Further, for the embodiment shown, the exemplary wasteheat recovery system 86 further includes an expansion device in fluid communication with thethermal transfer bus 98 downstream of theheat sink exchanger 96 and upstream of theheat source exchanger 250. The expansion device may be any suitable expansion device. For example, for the embodiment shown, the expansion device is configured as aturbine 114 in fluid communication with thethermal transfer bus 98 downstream of theheat sink exchanger 96 and upstream of theheat source exchanger 250. With such an embodiment, theturbine 114 may extract additional energy from the thermal transfer fluid, increasing an efficiency of the wasteheat recovery system 86 and thegas turbine engine 10. As will be appreciated, inclusion of the expansion device may generally allow for the reduction of a temperature of the thermal transfer fluid to a relatively low temperature such that the thermal transfer fluid may accept heat from the heat source through theheat source exchanger 250. For example, the expansion device may reduce a temperature of the thermal transfer fluid at least about one hundred degrees Celsius, such as at least about one hundred and fifty degrees Celsius, such as up to about one thousand degrees Celsius. - However, in other embodiments, the expansion device may not be configured to extract additional work form the thermal transfer fluid, and instead may simply be configured to expand the thermal transfer fluid (e.g., through an increasing cross-sectional area) to reduce a temperature and a pressure of the thermal transfer fluid. Further, although the expansion device/
turbine 114 is depicted schematically inward of thecore air flowpath 37 along a radial direction R of thegas turbine engine 10 ofFIG. 2 , in other embodiments, the expansion device/turbine 114 may instead be positioned outward of thecore air flowpath 37 along the radial direction R. - In one or more embodiments, the thermal transfer fluid may be a single phase thermal transfer fluid during operation of the waste
heat recovery system 86. In such a manner, the thermal transfer fluid may remain in substantially a liquid phase during operation. Alternatively, the thermal transfer fluid may be a phase change thermal transfer fluid during operation of the wasteheat recovery system 86. For example, the wasteheat recovery system 86 may generally operate on a refrigeration cycle, such that the thermal transfer fluid changes between liquid and gaseous phases during operation of the wasteheat recovery system 86. Additionally, or alternatively still, in one or more these configurations, the thermal transfer fluid may be in a supercritical phase during one or more stages of operation, or during all operations. For example, the thermal transfer fluid may be a supercritical carbon dioxide during certain operations or all operations of the wasteheat recovery system 86. - The exemplary
gas turbine engine 10 and wasteheat recovery system 86 depicted inFIG. 2 is, however, provided by way of example only. In other embodiments, the wasteheat recovery system 86 may have any other suitable configuration. As noted above, theheat sink exchanger 96 may be disposed in any other location that may use the heat from the wasteheat recovery system 86 to improve, e.g., efficiency. -
FIG. 3 is a schematic view of a portion of a turbine section and anexhaust section 32 of a gas turbine engine according to one or more embodiments. More specifically,FIG. 3 shows anLP turbine 30 and a heat source exchanger 250 disposed within theexhaust section 32 according to one or more embodiments. In certain exemplary embodiments, the turbine section andexhaust section 32 ofFIG. 3 may be configured in a similar manner as the exemplary turbine section andexhaust section 32 described above with reference toFIG. 1 . For example, the turbine section is generally configured as part of a gas turbine engine defining a radial direction R and an axial direction A. - The
LP turbine 30 according to one or more embodiments includes, in serial flow order and along the axial direction A, firstturbine guide vanes 210, first stage LPturbine rotor blades 221, secondturbine guide vanes 230, second stage LPturbine rotor blades 223, and final stage LPturbine rotor blades 241. Theexhaust section 32 according to one or more embodiments is disposed downstream of theLP turbine 30 in the axial direction A and includes theheat source exchanger 250 also downstream of the final stage LPturbine rotor blades 241 in the axial direction A. - The gas turbine engine further includes a waste
heat recovery system 86 with athermal transfer bus 98 having a thermal transfer fluid configured to flow therethrough. Theheat source exchanger 250 is part of the wasteheat recovery system 86 and is coupled to thethermal transfer bus 98, such that the thermal transfer fluid passing through thethermal transfer bus 98 passes through the heat source exchanger 250 to extract heat from theexhaust section 32 of thegas turbine engine 10. - According to one or more embodiments, the
heat source exchanger 250 is a frame-integrated heat exchanger that is integrated into the frame of theturbomachine 16. According to one or more embodiments, theheat source exchanger 250 is a heat exchanger formed separately from the frame of theturbomachine 16. - According to one or more embodiments, the first
turbine guide vanes 210 are directly upstream of the first stage LPturbine rotor blades 221, the first stage LPturbine rotor blades 221 are directly upstream of the secondturbine guide vanes 230, the secondturbine guide vanes 230 are directly upstream of the second stage LPturbine rotor blades 223, and the second stage LPturbine rotor blades 223 are directly upstream of the final stage LPturbine rotor blades 241, and the final stage LPturbine rotor blades 241 are directly upstream of theheat source exchanger 250. According to one or more embodiments, the firstturbine guide vanes 210 and the secondturbine guide vanes 230 are stationary. - While
FIGS. 1 and 3 show two stages of turbine rotor blades with a single stage of turbine guide vanes therebetween upstream of the final stage LPturbine rotor blades 241, one or more embodiments may include additional stages of turbine rotor blades and turbine guide vanes. For example, theLP turbine 30 may further include third turbine guide vanes and third stage LP turbine rotor blades in serial order downstream of the second stage LPturbine rotor blades 223 and may further include fourth turbine guide vanes and fourth stage LP turbine rotor blades in serial order downstream of the third stage LP turbine rotor blades, and so on. - The first stage LP
turbine rotor blades 221 and the second stage LPturbine rotor blades 223 rotate in a first direction around thelongitudinal centerline 12. The first stage LPturbine rotor blades 221 and the second stage LPturbine rotor blades 223 are connected to each other via a firstLP turbine spool 225 that is driven by the first stage LPturbine rotor blades 221 and the second stage LPturbine rotor blades 223 to rotate in the first direction around thelongitudinal centerline 12. The firstLP turbine spool 225 is coupled to agearbox 260. If more than two stages of LP turbine rotor blades are disposed upstream of the final stage LPturbine rotor blades 241, the additional stage(s) of LP turbine rotor blades are also connected to the first stage LPturbine rotor blades 221 and the second stage LPturbine rotor blades 223 via the firstLP turbine spool 225. - The final stage LP
turbine rotor blades 241 are connected to a secondLP turbine spool 245 and rotate in a second direction around thelongitudinal centerline 12, which is opposite the first direction. The final stage LPturbine rotor blades 241 are connected to the secondLP turbine spool 245 that is driven by the final stage LPturbine rotor blades 241 to rotate in the second direction. The secondLP turbine spool 245 is also coupled to thegearbox 260. - The
gearbox 260 is further coupled to the LP shaft orspool 36, such that the first stage LPturbine rotor blades 221, the second stage LPturbine rotor blades 223, and the final stage LPturbine rotor blades 241 drive the LP shaft orspool 36 through thegearbox 260. Thus, theLP turbine 30 supports operation of theLP compressor 22 and/or rotation of the fan 38 (seeFIG. 1 ). According to one or more embodiments, thegearbox 260 is disposed partially or entirely on an inner side of theLP turbine 30 in the radial direction R. - According to one or more embodiments, the
gearbox 260 may be structured as a planetary gear system in which that the LP shaft orspool 36 is fixed to or connected to asun gear 261, the secondLP turbine spool 245 is connected to planet gears 263 that are disposed around thesun gear 261 and are meshed with thesun gear 261, and the firstLP turbine spool 225 is fixed to or connected to aring gear 265 disposed around the planet gears 263 and is meshed with the planet gears 263. Thus, the firstLP turbine spool 225 and the LP shaft orspool 36 rotate in the first direction around thelongitudinal centerline 12, with the firstLP turbine spool 225 rotating at a higher speed than the LP shaft orspool 36, and the secondLP turbine spool 245 rotates in the second direction at a lower speed than the firstLP turbine spool 225. - As shown in
FIG. 3 , the final stage LPturbine rotor blades 241 have a significantly greater height in the radial direction R than the first stage LPturbine rotor blades 221 and the second stage LPturbine rotor blades 223. The final stage LPturbine rotor blades 241 also have a significantly greater height in the radial direction R than final stage LP turbine rotor blades of conventional high speed low pressure turbines that rotate at the same speed in the same direction as the preceding stages of LP turbine rotor blades. This greater height is possible due to the reduced speed of the final stage LPturbine rotor blades 241 rotating in the second direction, as the reduced speed reduces the stresses experienced by the final stage LPturbine rotor blades 241 compared to the aforementioned final stage LP turbine rotor blades of conventional high speed low pressure turbines. Furthermore, the reduction in stresses experienced by the final stage LPturbine rotor blades 241 may also enable materials that could not withstand the greater stresses experienced by the final stage LP turbine rotor blades of conventional high speed low pressure turbines. According to one or more embodiments, the first stage LPturbine rotor blades 221, the second stage LPturbine rotor blades 223, and the final stage LPturbine rotor blades 241 are formed of nickel alloys, or are formed of a material that comprises nickel alloys. According to one or more embodiments, the final stage LPturbine rotor blades 241 are formed of a different material from the first stage LPturbine rotor blades 221 and the second stage LPturbine rotor blades 223. According to one or more embodiments, the final stage LPturbine rotor blades 241 are formed of titanium aluminide or a material comprising titanium aluminide. As titanium aluminide is lighter than nickel alloys, forming the final stage LPturbine rotor blades 241 of titanium aluminide instead of nickel alloys, enabled by the lower speed of the final stage LPturbine rotor blades 241, results in significant weight savings that in turn enables a more efficientgas turbine engine 10. - According to one or more embodiments, the lower speed of rotation of the final stage LP
turbine rotor blades 241 also reduces a Mach number of the combustion gases exiting theLP turbine 30 and entering theexhaust section 32. For example, combustion gases may exit conventional high speed low pressure turbines and enter the exhaust section at a Mach number equal to around ½ Ma. According to one or more embodiments, theLP turbine 30 is structured such that combustion gases exit theLP turbine 30 and enter theexhaust section 32 at a Mach number of ⅓ Ma or lower. With a Mach number of ⅓ Ma or less, a frame-integrated heat exchanger that is integrated into the frame of theturbomachine 16 may be employed as the heat source exchanger 250 in theexhaust section 32. According to one or more embodiments, theLP turbine 30 is structured such that combustion gases exits theLP turbine 30 and enters theexhaust section 32 at a Mach number of ¼ Ma or less. With a Mach number of a ¼ Ma or lower, a traditional heat exchanger separate from the frame of theturbomachine 16 may be employed as the heat source exchanger 250 in theexhaust section 32. - Additionally, due to the counter-rotation between the final stage LP
turbine rotor blades 241 and the stage of the LP turbine rotor blades immediately upstream of the final stage LPturbine rotor blades 241, no guide vane is required therebetween. Thus, compared to a conventional high speed LP turbine in which the final stage LP turbine rotor blades rotate in the same direction as the other stages of LP turbine rotor blades, a guide vane can be removed, such that a more axiallycompact LP turbine 30 may be formed. Additionally, the more axiallycompact LP turbine 30 may in turn create more space in the axial direction A for theheat source exchanger 250. A largerheat source exchanger 250 extracts more heat from theexhaust section 32. - This written description uses examples, including the best mode, and also to enable any person skilled in the art to practice the disclosure, including making and using any devices or systems and performing any incorporated methods. The patentable scope is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal language of the claims.
- Further aspects are provided by the subject matter of the following clauses:
- A turbine section and an exhaust section for a gas turbine engine comprises a low pressure (LP) turbine comprising first stage LP turbine blades that rotate in a first direction at a first speed, and final stage LP turbine blades downstream of the first stage LP turbine blades that rotate in a second direction opposite the first direction at a second speed, wherein the second speed is lower than the first speed.
- The turbine section and the exhaust section of one or more of these clauses, wherein the LP turbine further comprises second stage LP turbine blades downstream of the first stage LP turbine blades and upstream of the final stage LP turbine blades that rotate in the first direction at the first speed, and guide vanes disposed between the first stage LP turbine blades and the second stage LP turbine blades.
- The turbine section and the exhaust section of one or more of these clauses, further comprising a heat source exchanger disposed downstream of the final stage LP turbine blades.
- The turbine section and the exhaust section of one or more of these clauses, wherein the first stage LP turbine blades are connected to a first LP turbine spool, wherein the final stage LP turbine blades are connected to a second LP turbine spool, and wherein the LP turbine further comprises a gearbox to which the first LP turbine spool and the second LP turbine spool are coupled.
- The turbine section and the exhaust section of one or more of these clauses, wherein the gearbox comprises a planetary gear system, wherein a ring gear of the planetary gear system is connected to the first LP turbine spool, wherein planet gears of the planetary gear system are connected to the second LP turbine spool, and wherein a sun gear of the planetary gear system is configured to be connected to an LP shaft or a spool that is connected to a fan of the gas turbine engine.
- The turbine section and the exhaust section of one or more of these clauses, wherein the last stage LP turbine blades are formed of a different material from the first stage LP turbine blades.
- The turbine section and the exhaust section of one or more of these clauses, wherein the final stage LP turbine blades are the only turbine rotor blades of the LP turbine configured to rotate in the second direction.
- The turbine section and the exhaust section of one or more of these clauses, further comprising a high pressure (HP) turbine located upstream of the LP turbine.
- A turbine section and an exhaust section of a gas turbine engine comprises a low pressure (LP) turbine comprising a plurality of stages of LP turbine blades that rotate in a first direction at a first speed, and final stage LP turbine blades downstream of the plurality of stages of LP turbine blades that rotate in a second direction opposite the first direction at a second speed, wherein the second speed is lower than the first speed.
- The turbine section and the exhaust section of one or more of these clauses, wherein the LP turbine further comprises a drum positioned inward of an airflow path through the LP turbine along a radial direction of the gas turbine engine, wherein the plurality of stages of LP turbine blades that rotate in the first direction are coupled together with the drum, wherein each of the final stages LP turbine blades includes an outer end along the radial direction, and wherein each of the final stages LP turbine blades is mechanically isolated at the outer end.
- The turbine section and the exhaust section of one or more of these clauses, wherein the LP turbine further comprises guide vanes disposed between adjacent stages of the plurality of stages of LP turbine blades.
- The turbine section and the exhaust section of one or more of these clauses, wherein the final stage LP turbine blades are disposed directly downstream of a furthest downstream stage of LP turbine blades of the plurality of stages of LP turbine blades.
- The turbine section and the exhaust section of one or more of these clauses, further comprising a heat source exchanger disposed downstream of the final stage LP turbine blades.
- The turbine section and the exhaust section of one or more of these clauses, wherein each of the plurality of stages of LP turbine blades is connected to a first LP turbine spool, wherein the final stage LP turbine blades are connected to a second LP turbine spool, and wherein the LP turbine further comprises a gearbox to which the first LP turbine spool and the second LP turbine spool are coupled.
- The turbine section and the exhaust section of one or more of these clauses, wherein the gearbox comprises a planetary gear system, wherein a ring gear of the planetary gear system is connected to the first LP turbine spool, wherein planet gears of the planetary gear system are connected to the second LP turbine spool, and wherein a sun gear of the planetary gear system is configured to be connected to a LP shaft or spool that is connected to a fan of the gas turbine engine.
- A gas turbine engine comprises a low pressure (LP) turbine comprising a plurality of stages of LP turbine blades that rotate in a first direction at a first speed, and final stage LP turbine blades downstream of the plurality of stages of LP turbine blades that rotate in a second direction opposite the first direction at a second speed, wherein the second speed is lower than the first speed.
- The gas turbine engine of one or more of these clauses, wherein each of the plurality of stages of LP turbine blades is connected to a first LP turbine spool, wherein the final stage LP turbine blades are connected to a second LP turbine spool, and wherein the LP turbine further comprises a gearbox to which the first LP turbine spool and the second LP turbine spool are coupled.
- The gas turbine engine of one or more of these clauses, wherein the gearbox comprises a planetary gear system, wherein a ring gear of the planetary gear system is connected to the first LP turbine spool, wherein planet gears of the planetary gear system are connected to the second LP turbine spool, and wherein a sun gear of the planetary gear system is connected to an LP shaft or a spool that is connected to a fan of the gas turbine engine.
- The gas turbine engine of one or more of these clauses, wherein the LP turbine is structured such that, during operation of the gas turbine engine, combustion gases exiting the LP turbine do not exceed a Mach number of ⅓ Ma.
- The gas turbine engine of one or more of these clauses, further comprising a heat recovery system, wherein the heat recovery system comprises a heat source exchanger disposed downstream of the final stage LP turbine blades, and a heat sink exchanger disposed in a portion of the gas turbine engine upstream of a combustion section of the gas turbine engine.
Claims (20)
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